With reference to FIG. 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, and a core engine including an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the core engine and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
Large ducted fan gas turbine engines are conventionally fitted with a thrust reverser which reverses the second air flow B on landing to assist with aircraft deceleration. The thrust reverser typically comprises a series of fixed cascade boxes which contain turning vanes for channelling airflow from the bypass duct in a forward direction, the cascade boxes being installed in the nacelle 21 around the bypass duct 22. The nacelle is formed with a pair of translating sleeve cowls which, in normal operation, shield the cascade boxes from the external airflow and from the second air flow B. To operate the thrust reverser, the sleeve cowls are translated rearwardly to expose the cascade boxes both internally and externally and to deploy blocker doors which block the bypass duct 22 and force the second air flow B through the cascade boxes. Conventionally, the blocker doors are deployed via drag links which are each attached at one end to the respective door and at the other end to a fixed structure at the inner surface of the bypass duct 22
In FIG. 1, the bypass exhaust nozzle 23 has a fixed area. However, variable area bypass nozzles are known and can be used to control the discharge area for the second air flow B so that the engine can operate efficiently over a wider range of operating conditions. In particular, a present trend towards lower pressure ratio engines increases the desirability of controlling fan stability margins by varying the bypass exhaust nozzle discharge area.
Combining elements of a thrust reverser and a variable area nozzle can help to reduce engine weight, as well as providing other operational advantages.
U.S. Pat. No. 5,655,360 proposes a thrust reverser for both modulating and reversing bypass flow discharged from a fan through a bypass duct. The reverser includes an aft cowl which has an aft end surrounding the core engine to define a discharge nozzle, and is translatable rearwardly to increase the discharge area of that nozzle. Further rearward translation, however, exposes cascade turning vanes and causes deployment of blocker doors to produce reverse thrust. A disadvantage, however, of the proposal of U.S. Pat. No. 5,655,360 is that to produce large increases in the discharge area of the nozzle it is necessary for the aft cowl to be translated a relatively long distance. This complicates both the actuation mechanism for the cowl translation, and also the actuation mechanism for deployment of the blocker doors which must only be deployed when further rearward translation exposes the cascade turning vanes. A further disadvantage of this arrangement is that a single actuator varies the area of the discharge nozzle, and provides thrust reversal. Such a system poses a risk that the thrust reverser could deploy in flight, which may lead to the loss of the aircraft. Complex safety systems would therefore be required, further complicating the design. US 2008/0010969 proposes a nacelle for a gas turbine engine having a first cowl and a second cowl which is repositionable with respect to the first cowl. The second cowl can be positioned to provide either external wing blowing in a first position (to be used for example on landing) or thrust reversal in a second position. In the first position, the total quantity of airflow channelled through the fan nozzle is reduced to provide reduced thrust through the nozzle when external blowing is provided. The total effective flow area of the gas turbine is thereby either maintained or reduced in the first position in comparison to the second position.